The present invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in turbine stages which power the compressor and perform work by powering a fan in a typical turbofan aircraft engine application.
Turbine efficiency is maximized by maximizing inlet temperature of the combustion gases thereto, but the various turbine components must be suitably cooled for promoting long useful life.
Typical turbine components include stator vanes in the turbine nozzles which channel the combustion gases to a corresponding row of turbine rotor blade extending radially outwardly from a supporting rotor disk.
The stator vanes and rotor blades have typical airfoil configurations specifically configured for maximizing energy extraction from the hot combustion gases. The airfoils are typically hollow and include internal cooling circuits through which pressurized air bled from the compressor is used as a coolant for internally cooling the airfoils during operation.
The spent cooling air is typically discharged through various rows of film cooling holes which extend through the sidewalls of the airfoil for creating thin films of cooling air over the outer surface of the airfoils to provide thermal insulation.
Any air bled from the compressor which is not used in the combustion cycle decreases the overall efficiency of the engine and therefore is typically minimized. However, the stator vanes and rotor blades have complex 3D airfoil configurations starting at the leading edges thereof and increasing in thickness to a maximum width hump region behind the leading edge, with the airfoil then tapering to a narrow and thin trailing edge.
The aerodynamic performance of the airfoils effects corresponding distributions of velocity, pressure, and heat over the generally concave pressure sides and the generally convex suction sides axially between the opposite leading and trailing edges and radially between the radially inner root and outer tip.
The internal cooling circuits of the airfoils are specifically configured to match the external flow environment and tailor use of the limited coolant bled from the compressor.
The prior art is replete with numerous patents dating back over decades which disclose various configurations for cooling the different parts of turbine airfoils with corresponding benefits, as well as disadvantages.
Adding to the complexity of modern turbine airfoil cooling design is the size and manufacture of these components. Large engines have large turbine airfoils and require correspondingly large coolant flow for effective cooling.
Small engines have correspondingly small turbine airfoils which nevertheless require suitable cooling since the combustion gas temperature is as high as possible for maximizing engine efficiency irrespective of engine size.
One significant problem in manufacturing small turbine airfoils is the ability to cast small airfoils with correspondingly small cooling features.
The modern turbine airfoil is typically cast from a nickel based superalloy metal having enhanced strength at high operating temperature. Superalloy airfoils are typically made by casting, which requires ceramic cores that define the small internal cooling features of the airfoil.
However, small ceramic cores are particularly brittle and subject to damage, and therefore increased waste during manufacture. And, cooling features incorporated into the cores have minimal practical sizes which may nevertheless be excessive for the small engine environment.
More specifically, a typical turbine airfoil may include internal impingement holes in corresponding partitions through which the coolant is ejected in small jets for internally impingement cooling various portions of the airfoil.
Correspondingly, the turbine airfoil typically also includes various rows of film cooling holes through the sidewalls thereof that discharge the spent internal cooling air to create the external air thermal insulation.
Whereas the external film cooling holes may be readily formed after airfoil casting using conventional drilling techniques, the internal impingement holes must be formed by casting since internal access is unavailable for drilling.
The significance of these differences is that drilled film cooling holes may be made substantially smaller in diameter than internal cast impingement holes.
The typical film cooling hole is cylindrical in cross section and may have a diameter of about 12-18 mils (0.3-0.46 mm). Such small holes can readily be drilled, but are not typically castable.
Typical internal impingement holes also have cylindrical cross sections and may range in diameter from about 24-90 mils (0.6-2.2 mm) as a function of turbine airfoil size from small to large.
However, the minimum impingement hole size is typically limited by the smallest practical casting size, yet that minimum casting size results in a relatively large impingement hole in small turbine airfoils.
Accordingly, excess coolant will be channeled through even the smallest castable impingement hole in a small turbine airfoil and undesirably reduce the overall efficiency of the engine.
Accordingly, it is desired to provide a turbine airfoil having improved cooling features, and in particular useful for small turbine airfoils.